U.S. Department

of Transportation

Federal Aviation

Administration

Advisory

Circular

Subject: AEROELASTIC STABILITY Date: Draft AC No: 25.629-1X

 SUBSTANTIATION OF TRANSPORT Initiated by: ANM-115

 CATEGORY AIRPLANES

1. PURPOSE. This advisory circular (AC) provides guidance material for 

acceptable means, but not the only means, of demonstrating compliance with the 

provisions of Title 14, Code of Federal Regulations (14 CFR), part 25 associated with the 

design requirements for transport category airplanes to preclude the aeroelastic 

instabilities of flutter, divergence, and control reversal. The precise details for analytical 

procedures and testing techniques are beyond the scope of this AC. However, some 

general information is included, with supporting discussions, for consideration when 

demonstrating compliance with § 25.629 and related regulations.

2. CANCELLATION. This AC cancels Advisory Circular 25.629-1A, Aeroelastic 

Stability Substantiation of Transport Category Airplanes, dated 7/23/98.

3. RELATED CFR SECTIONS.

§ 25.251 - Vibration and buffeting.

§ 25.305 - Strength and deformation.

§ 25.335 - Design airspeeds.

§ 25.343 - Design fuel and oil loads.

§ 25.571 - Damage-tolerance and fatigue evaluation of structure.

§ 25.629 - Aeroelastic stability requirements.

§ 25.631 - Bird strike damage.

§ 25.671 - General (Control Systems).

 § 25.672 - Stability augmentation and automatic and power-operated systems.

 § 25.1309 - Equipment, systems and installations.

 § 25.1329 - Automatic pilot system.

 § 25.1419 - Ice protection.
DATE AC 25.629-1X

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4. BACKGROUND.

a. Flutter and other aeroelastic instability phenomena have had a significant 

influence on airplane development and the airworthiness criteria governing the design of 

civil airplanes. The initial requirement for consideration of flutter was minimal in the

1931 "Airworthiness Requirements of Air Commercial Regulations for Aircraft," Bulletin 

No. 7-A. The airplane flutter requirement specified that "no surface shall show any signs 

of flutter or appreciable vibration in any attitude or condition of flight." In 1934 Bulletin 

No. 7-A was revised in view of service experience and contained advice and good 

practice techniques for the early airplane designer regarding flutter prevention measures. 

All airplane designs were required to have interconnected elevators, statically-balanced 

ailerons, irreversible or balanced tabs, and, in some cases, a ground vibration test was 

required to be conducted.

b. Regulations dealing specifically with flutter, deformation, and vibration on 

transport category airplanes were first introduced when part 04 of the Civil Air 

Regulations (CAR) became effective in the mid-1940s. The criteria related the solution 

of the flutter problem to frequency ratios based on model tests conducted by the Army Air 

Corps. Also, based on the Army Air Corps developments, part 04 imposed a design 

factor of 1.2 on equivalent airspeed to provide a stiffness margin for the airframe. In 

addition to this empirical approach, and recognizing the advancing state-of-the-art, part 

04 referenced publications containing developing flutter theory.

c. The flutter requirement of part 04 evolved into CAR 4b.308 where 

developing fail-safe philosophy continued to change the scope of flutter substantiation. 

Among these developments was a revision to CAR 4b.320 in 1956 to require fail-safe 

tabs and a revision to CAR 4b.308 in 1959 to require fail-safe flutter damper installations. 

The flutter requirement was extensively revised in 1964 to require compliance with the 

single failure criteria for the entire airplane as well as adding special provisions for 

turboprop airplanes.

d. Service experience indicated that single failure criteria related to flutter 

stability were not sufficiently objective and comprehensive to cover modern, complex, 

transport airplanes equipped with highly redundant systems. Therefore, part 25, which 

was recodified from part 04b of the CAR, was amended to require that, unless 

combinations of failures are shown to be extremely improbable, they must be considered 

in design for freedom from flutter and divergence.

e. The development of speed and attitude limiting systems has created the 

need for a minimum speed margin for fail-safe aeroelastic stability substantiation. 

Part 25 as amended at Amendment 25-77 incorporated this minimum fail-safe speed 

boundary, revised the safety margins for aeroelastic stability, and expanded the list of 

failures, malfunctions and adverse conditions that needed to be addressed. 
DATE AC 25.629-1X

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f. Additional regulations governing the interaction of systems with structures 

have been written for airplanes with advanced electronic flight control systems. These 

regulations prescribe variations in the fail-safe speed margins depending on the 

probability of system failure.

5. DISCUSSION OF REQUIREMENTS. The general requirement for 

demonstrating freedom from aeroelastic instability is contained in § 25.629, which also 

sets forth specific requirements for the investigation of these aeroelastic phenomena for 

various airplane configurations and flight conditions. Additionally, there are other 

conditions defined by the sections of part 25 listed in paragraph 3, above, to be 

investigated for aeroelastic stability to assure safe flight. Many of the conditions 

contained in this AC pertain only to certain specific amendments of part 25. Type design 

changes to airplanes certified to an earlier part 25 amendment must meet the certification 

basis established for the modified airplane.

a. Aeroelastic Stability Envelope.

(1) For nominal conditions without failures, malfunctions, or adverse 

conditions, freedom from aeroelastic instability is required to be shown for all 

combinations of airspeed and altitude encompassed by the design dive speed (VD) and 

design dive Mach number (MD) versus altitude envelope enlarged at all points by an 

increase of 15 percent in equivalent airspeed at both constant Mach number and constant 

altitude. Figure 1A represents a typical design envelope expanded to the required 

aeroelastic stability envelope. Note that some required Mach number and airspeed 

combinations correspond to altitudes below standard sea level.

(2) The aeroelastic stability envelope may be limited to a maximum 

Mach number of 1.0 when MD is less than 1.0 and when there is no large and rapid 

reduction in damping as MD is approached.

(3) Some configurations and conditions that are required to be 

investigated by § 25.629 and other part 25 regulations consist of failures, malfunctions or 

adverse conditions. Aeroelastic stability investigations of these fail-safe conditions need 

to be carried out for all approved altitudes to the greater airspeed defined by:

(a) The VD/MD envelope determined by § 25.335(b); or,

(b) An altitude-airspeed envelope defined by a 15 percent 

increase in equivalent airspeed above VC at constant altitude, from sea level up to the 

altitude of the intersection of 1.15 VC with the extension of the constant cruise Mach 

number line, MC, then a linear variation in equivalent airspeed to MC + 0.05 at the altitude 

of the lowest VC/MC intersection; then at higher altitudes, up to the maximum flight 

altitude, the boundary defined by a 0.05 Mach increase in MC at constant altitude.
DATE AC 25.629-1X

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Figure 1B shows the minimum aeroelastic stability envelope for fail-safe 

conditions, which is a composite of the highest speed at each altitude from either the VD

envelope or the constructed altitude-airspeed envelope based on the defined VC and MC.

Fail-safe design speeds, other than the ones defined above, may be used for 

certain system failure conditions when specifically authorized by other rules or special 

conditions prescribed in the certification basis of the airplane.

FIGURE 1A. MINIMUM REQUIRED AEROELASTIC STABILITY MARGIN

FIGURE 1B. MINIMUM FAIL-SAFE CLEARANCE ENVELOPE
DATE AC 25.629-1X

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b. Configurations and Conditions. The following paragraphs provide a 

summary of the configurations and conditions to be investigated in demonstrating 

compliance with part 25. Specific design configurations may warrant additional 

considerations not discussed in this AC.

(1) Nominal Configurations and Conditions. Nominal configurations 

and conditions of the airplane are those that are likely to exist during normal operation. 

Freedom from aeroelastic instability should be shown throughout the expanded clearance 

envelope described in paragraph 5a(1) above for:

(a) The range of fuel and payload combinations, including zero 

fuel, for which certification is requested.

(b) Configurations with ice mass accumulations on unprotected 

surfaces for airplanes approved for operation in icing conditions. See paragraph 6a(4)(e) 

of this AC.

(c) All normal combinations of autopilot, yaw damper, or other 

automatic flight control systems.

(d) All possible engine settings and combinations of settings 

from idle power to maximum available thrust including the conditions of one engine 

stopped and windmilling, in order to address the influence of gyroscopic loads and thrust 

on aeroelastic stability.

(2) Failures, Malfunctions, and Adverse Conditions. The following 

conditions should be investigated for aeroelastic instability within the fail-safe envelope 

defined in paragraph 5a(3) above.

(a) Any critical fuel loading conditions, not shown to be 

extremely improbable, which may result from mismanagement of fuel.

(b) Any single failure in any flutter control system.

(c) For airplanes not approved for operation in icing 

conditions, ice accumulation expected as a result of an inadvertent encounter. For 

airplanes approved for operation in icing conditions, the ice accumulation expected as the 

result of any single failure in the de-icing system, or any combination of failures not 

shown to be extremely improbable. See paragraph 6a(4)(e) of this AC.

(d) Failure of any single element of the structure supporting 

any engine, independently mounted propeller shaft, large auxiliary power unit, or large 

externally mounted aerodynamic body (such as an external fuel tank).
DATE AC 25.629-1X

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(e) For airplanes with engines that have propellers or large 

rotating devices capable of significant dynamic forces, any single failure of the engine 

structure that would reduce the rigidity of the rotational axis.

(f) The absence of aerodynamic or gyroscopic forces resulting 

from the most adverse combination of feathered propellers or other rotating devices 

capable of significant dynamic forces. In addition, the effect of a single feathered 

propeller or rotating device must be coupled with the failures of paragraphs 5b(2)(d) and 

5b(2)(e) above.

(g) Any single propeller or rotating device capable of 

significant dynamic forces rotating at the highest likely overspeed.

(h) Any damage or failure condition, required or selected for 

investigation by § 25.571. The single structural failures described in paragraphs 5b(2)(d) 

and 5b(2)(e) above need not be considered in showing compliance with this paragraph if;

1 The structural element could not fail due to discrete 

source damage resulting from the conditions described in § 25.571(e); and

2 A damage tolerance investigation in accordance 

with § 25.571(b) shows that the maximum extent of damage assumed for the purpose of 

residual strength evaluation does not involve complete failure of the structural element.

(i) Any damage, failure or malfunction, considered under 

§§ 25.631, 25.671, 25.672, and 25.1309. This includes the condition of two or more 

engines stopped or windmilling for the design range of fuel and payload combinations, 

including zero fuel.

(j) Any other combination of failures, malfunctions, or adverse 

conditions not shown to be extremely improbable.

c. Detail Design Requirements.

(1) Main surfaces, such as wings and stabilizers, should be designed to 

meet the aeroelastic stability criteria for nominal conditions and should be investigated 

for meeting fail-safe criteria by considering stiffness changes due to discrete damage or 

by reasonable parametric variations of design values.

(2) Control surfaces, including tabs, should be investigated for 

nominal conditions and for failure modes that include single structural failures (such as 

actuator disconnects, hinge failures, or, in the case of aerodynamic balance panels, failed 

seals), single and dual hydraulic system failures and any other combination of failures not 

shown to be extremely improbable. Where other structural components contribute to the 
DATE AC 25.629-1X

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aeroelastic stability of the system, failures of those components should be considered for 

possible adverse effects.

(3) Where aeroelastic stability relies on control system stiffness and/or 

damping, additional conditions should be considered. The actuation system should 

continuously provide, at least, the minimum stiffness or damping required for showing 

aeroelastic stability without regard to probability of occurrence for:

(a) More than one engine stopped or windmilling,

(b) Any discrete single failure resulting in a change of the 

structural modes of vibration (for example; a disconnect or failure of a mechanical 

element, or a structural failure of a hydraulic element, such as a hydraulic line, an 

actuator, a spool housing or a valve),

(c) Any damage or failure conditions considered under 

§§ 25.571, 25.631 and 25.671.

The actuation system minimum requirements should also be continuously met 

after any combination of failures not shown to be extremely improbable (occurrence less 

than 10-9 per flight hour). However, certain combinations of failures, such as dual 

electric or dual hydraulic system failures, or any single failure in combination with any 

probable electric or hydraulic system failure (§ 25.671), are not normally considered 

extremely improbable regardless of probability calculations. The reliability assessment 

should be part of the substantiation documentation. In practice, meeting the above 

conditions may involve design concepts such as the use of check valves and 

accumulators, computerized pre-flight system checks and shortened inspection intervals 

to protect against undetected failures.

(4) Consideration of free play may be incorporated as a variation in 

stiffness to assure adequate limits are established for wear of components such as control 

surface actuators, hinge bearings, and engine mounts in order to maintain aeroelastic 

stability margins.

(5) If balance weights are used on control surfaces, their effectiveness 

and strength, including that of their support structure, should be substantiated.

(6) The automatic flight control system should not interact with the 

airframe to produce an aeroelastic instability. When analyses indicate possible adverse 

coupling, tests should be performed to determine the dynamic characteristics of actuation 

systems such as servo-boost, fully powered servo-control systems, closed-loop airplane 

flight control systems, stability augmentation systems, and other related powered-control 

systems.
DATE AC 25.629-1X

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6. COMPLIANCE. Demonstration of compliance with aeroelastic stability 

requirements for an airplane configuration may be shown by analyses, tests, or some 

combination thereof. In most instances, analyses are required to determine aeroelastic 

stability margins for normal operations, as well as for possible failure conditions. Wind 

tunnel flutter model tests, where applicable, may be used to supplement flutter analyses. 

Ground testing may be used to collect stiffness or modal data for the airplane or 

components. Flight testing may be used to demonstrate compliance of the airplane design 

throughout the design speed envelope.

a. Analytical Investigations. Analyses should normally be used to 

investigate the aeroelastic stability of the airplane throughout its design flight envelope 

and as expanded by the required speed margins. Analyses are used to evaluate aeroelastic 

stability sensitive parameters such as aerodynamic coefficients, stiffness and mass 

distributions, control surface balance requirements, fuel management schedules, 

engine/store locations, and control system characteristics. The sensitivity of most critical 

parameters may be determined analytically by varying the parameters from nominal. 

These investigations are an effective way to account for the operating conditions and 

possible failure modes which may have an effect on aeroelastic stability margins, and to 

account for uncertainties in the values of parameters and expected variations due to inservice wear or failure conditions.

(1) Analytical Modeling. The following sections discuss acceptable, 

but not the only, methods and forms of modeling airplane configurations and/or 

components for purposes of aeroelastic stability analysis. The types of investigations 

generally encountered in the course of airplane aeroelastic stability substantiation are also 

discussed. The basic elements to be modelled in aeroelastic stability analyses are the 

elastic, inertial, and aerodynamic characteristics of the system. The degree of complexity 

required in the modeling, and the degree to which other characteristics need to be 

included in the modeling, depend upon the system complexity.

(a) Structural Modeling. Most forms of structural modeling 

can be classified into two main categories: (1) modeling using a lumped mass beam, and 

(2) finite element modeling. Regardless of the approach taken for structural modeling, a 

minimum acceptable level of sophistication, consistent with configuration complexity, is 

necessary to satisfactorily represent the critical modes of deformation of the primary 

structure and control surfaces. The model should reflect the support structure for the 

attachment of control surface actuators, flutter dampers, and any other elements for which 

stiffness is important in prevention of aeroelastic instability. Wing-pylon mounted 

engines are often significant to aeroelastic stability and warrant particular attention in the 

modeling of the pylon, and pylon-engine and pylon-wing interfaces. The model should 

include the effects of cut-outs, doors, and other structural features which may tend to 

affect the resulting structural effectiveness. Reduced stiffness should be considered in the 

modeling of airplane structural components which may exhibit some change in stiffness 

under limit design flight conditions. Structural models include mass distributions as well 

as representations of stiffness and possibly damping characteristics. Results from the 
DATE AC 25.629-1X

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models should be compared to test data, such as that obtained from ground vibration 

tests, in order to determine the accuracy of the model and its applicability to the 

aeroelastic stability investigation.

(b) Aerodynamic Modeling.

1 Aerodynamic modeling for aeroelastic stability 

requires the use of unsteady, two-dimensional strip or three-dimensional panel theory 

methods for incompressible or compressible flow. The choice of the appropriate 

technique depends on the complexity of the dynamic structural motion of the surfaces 

under investigation and the flight speed envelope of the airplane. Aerodynamic modeling 

should be supported by tests or previous experience with applications to similar 

configurations.

2 Main and control surface aerodynamic data are 

commonly adjusted by weighting factors in the aeroelastic stability solutions. The 

weighting factors for steady flow (k=0) are usually obtained by comparing wind tunnel 

test results with theoretical data. Special attention should be given to control surface 

aerodynamics because viscous and other effects may require more extensive adjustments 

to theoretical coefficients. Main surface aerodynamic loading due to control surface 

deflection should be considered.

(2) Types of Analyses.

(a) Oscillatory (flutter) and non-oscillatory (divergence and 

control reversal) aeroelastic instabilities should be analyzed to show compliance with 

§ 25.629.

(b) The flutter analysis methods most extensively used involve 

the modal analysis with unsteady aerodynamic forces derived from various two- and 

three-dimensional theories. These methods are generally for linear systems. Analyses 

involving control system characteristics should include equations describing system 

control laws in addition to the equations describing the structural modes.

(c) Airplane lifting surface divergence analyses should include 

all appropriate rigid body mode degrees-of-freedom since divergence may occur for a 

structural mode or the short period mode.

(d) Loss of control effectiveness (control reversal) due to the 

effects of elastic deformations should be investigated. Analyses should include the 

inertial, elastic, and aerodynamic forces resulting from a control surface deflection.
DATE AC 25.629-1X

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(3) Damping Requirements.

(a) There is no intent in this AC to define a flight test level of 

acceptable minimum damping.

(b) Flutter analyses results are usually presented graphically in 

the form of frequency versus velocity (V-f, Figure 2) and damping versus velocity (V-g, 

Figures 3 and 4) curves for each root of the flutter solution.

(c) Figure 3 details one common method for showing 

compliance with the requirement for a proper margin of damping. It is based on the 

assumption that the structural damping available is 0.03 (1.5% critical viscous damping) 

and is the same for all modes as depicted by the V-g curves shown in Figure 3. No 

significant mode, such as curves (2) or (4), should cross the g=0 line below VD or the 

g=0.03 line below 1.15 VD. An exception may be a mode exhibiting damping 

characteristics similar to curve (1) in Figure 3, which is not critical for flutter. A 

divergence mode, as illustrated by curve (3) where the frequency approaches zero, should 

have a divergence velocity not less than 1.15 VD.

(d) Figure 4 shows another common method of presenting the 

flutter analysis results and defining the structural damping requirements. An appropriate 

amount of structural damping for each mode is entered into the analysis prior to the flutter 

solution. The amount of structural damping used should be supported by measurements 

taken during full scale tests. This results in modes offset from the g=0 line at zero 

airspeed and, in some cases, flutter solutions different from those obtained with no 

structural damping. The similarity in the curves of Figures 3 and 4 are only for 

simplifying this example. The minimum acceptable damping line applied to the 

analytical results as shown in Figure 4 corresponds to 0.03 or the modal damping 

available at zero airspeed for the particular mode of interest, whichever is less, but in no 

case less than 0.02. No significant mode should cross this line below VD or the g=0 line 

below 1.15 VD.

(e) For analysis of failures, malfunctions or adverse conditions 

being investigated, the minimum acceptable damping level obtained analytically would 

be determined by use of either method above, but with a substitution of VC for VD and the 

fail-safe envelope speed at the analysis altitude as determined by 5a(3) above.
DATE AC 25.629-1X

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FIGURE 2. FREQUENCY VERSUS VELOCITY

FIGURE 3. DAMPING VERSUS VELOCITY - Method 1
DATE AC 25.629-1X

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FIGURE 4. DAMPING VERSUS VELOCITY - Method 2

(4) Analysis Considerations. Airframe aeroelastic stability analyses 

may be used to verify the design with respect to the structural stiffness, mass, fuel 

(including in-flight fuel management), automatic flight control system characteristics, and 

altitude and Mach number variations within the design flight envelope. The complete 

airplane should be considered as composed of lifting surfaces and bodies, including all 

primary control surfaces which can interact with the lifting surfaces to affect flutter 

stability. Control surface flutter can occur in any speed regime and has historically been 

the most common form of flutter. Lifting surface flutter is more likely to occur at high 

dynamic pressure and at high subsonic and transonic Mach numbers. Analyses are 

necessary to establish the mass balance and/or stiffness and redundancy requirements for 

the control surfaces and supporting structure and to determine the basic surface flutter 

trends. The analyses may be used to determine the sensitivity of the nominal airplane 

design to aerodynamic, mass, and stiffness variations. Sources of stiffness variation may 

include the effects of skin buckling at limit load factor, air entrapment in hydraulic 

actuators, expected levels of in-service free play, and control system components which 

may include elements with nonlinear stiffness. Mass variations include the effects of fuel 

density and distribution, control surface repairs and painting, and water and ice 

accumulation. 

(a) Control Surfaces. Control surface aeroelastic stability 

analyses should include control surface rotation, tab rotation (if applicable), significant 

modes of the airplane, control surface torsional degrees-of-freedom, and control surface 

bending (if applicable). Analyses of airplanes with tabs should include tab rotation that is 

both independent and related to the parent control surface. Control surface rotation 

frequencies should be varied about nominal values as appropriate for the condition. The 

control surfaces should be analysed as completely free in rotation unless it can be shown 

that this condition is extremely improbable. All conditions between stick-free and stickfixed should be investigated. Freeplay effects should be incorporated to account for any 
DATE AC 25.629-1X

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influence of in-service wear on flutter margins. The aerodynamic coefficients of the 

control surface and tab used in the aeroelastic stability analysis should be adjusted to 

match experimental values at zero frequency. Once the analysis has been conducted with 

the nominal, experimentally adjusted values of hinge moment coefficients, the analysis 

should be conducted with parametric variations of these coefficients and other parameters 

subject to variability. If aeroelastic stability margins are found to be sensitive to these 

parameters, then additional verification in the form of model or flight tests may be 

required.

(b) Mass Balance.

1 The magnitude and spanwise location of control 

surface balance weights may be evaluated by analysis and/or wind tunnel flutter model 

tests. If the control surface torsional degrees of freedom are not included in the analysis, 

then adequate separation needs to be maintained between the frequency of the control 

surface first torsion mode and the flutter mode.

2 Control surface unbalance tolerances should be 

specified to provide for repair and painting. The accumulation of water, ice, and/or dirt in 

or near the trailing edge of a control surface should be avoided. Free play between the 

balance weight, the support arm, and the control surface should not be allowed. Control 

surface mass properties (weight and static unbalance) should be confirmed by 

measurement before ground vibration testing.

3 The balance weights and their supporting structure 

should be substantiated for the extreme load factors expected throughout the design flight 

envelope. In the absence of a rational investigation, the following limit accelerations, 

applied through the balance weight center of gravity should be used.

100g normal to the plane of the surface

30g parallel to the hinge line

30g parallel to the plane of the surface and perpendicular to the hinge line

(c) Passive Flutter Dampers. Control surface passive flutter 

dampers may be used to prevent flutter in the event of failure of some element of the 

control surface actuation system or to prevent control surface buzz. Flutter analyses 

and/or flutter model wind tunnel tests may be used to verify adequate damping. Damper 

support structure flexibility should be included in the determination of adequacy of 

damping at the flutter frequencies. Any single damper failure should be considered. 

Combinations of multiple damper failures should be examined when not shown to be 

extremely improbable. The combined free play of the damper and supporting elements 

between the control surface and fixed surfaces should be considered. Provisions for inservice checks of damper integrity should be considered. Refer to paragraph 5c(3) above 

for conditions to consider where a control surface actuator is switched to the role of an 

active or passive damping element of the flight control system.
DATE AC 25.629-1X

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(d) Intersecting Lifting Surfaces. Intersecting lifting surface 

aeroelastic stability characteristics are more difficult to predict accurately than the 

characteristics of planar surfaces such as wings. This is due to difficulties both in 

correctly predicting vibration modal characteristics and in assessing those aerodynamic 

effects which may be of second order importance on planar surfaces, but are significant 

for intersecting surfaces. Proper representation of modal deflections and unsteady 

aerodynamic coupling terms between surfaces is essential in assessing the aeroelastic 

stability characteristics. The in-plane forces and motions of one or the other of the 

intersecting surfaces may have a strong effect on aeroelastic stability; therefore, the 

analysis should include the effects of steady flight forces and elastic deformations on the 

in-plane effects.

(e) Ice Accumulation. Aeroelastic stability analyses should 

use the mass distributions derived from ice accumulation up to and including those that 

can accrete in the applicable icing conditions in part 25, Appendices C and O. This 

includes any accretions that could develop on control surfaces. The analyses need not 

consider the aerodynamic effects of ice shapes. For airplanes approved for operation in 

icing conditions, all of the part 25, Appendix C, icing conditions and the Appendix O

icing conditions for which certification is sought are applicable. For airplanes exempt 

from § 25.1420, no evaluation of Appendix O icing conditions is required. For airplanes 

not approved for operation in icing conditions, all of the Appendix C and O icing 

conditions are applicable since the inadvertent encounter discussed in paragraph 5b(2)(c) 

of this AC can occur in any icing condition. For all airplanes, the ice accumulation 

determination should take into account the ability to detect the ice and, if appropriate, the 

time required to leave the icing condition.

(f) Whirl Flutter.

1 The evaluation of the aeroelastic stability should 

include investigations of any significant elastic, inertial, and aerodynamic forces, 

including those associated with rotations and displacements in the plane of any turbofan 

or propeller, including propeller or fan blade aerodynamics, powerplant flexibilities, 

powerplant mounting characteristics, and gyroscopic coupling.

2 Failure conditions are usually significant for whirl 

instabilities. Engine mount, engine gear box support, or shaft failures which result in a 

node line shift for propeller hub pitching or yawing motion are especially significant.

3 A wind tunnel test with a component flutter model, 

representing the engine/propeller system and its support system along with correlative 

vibration and flutter analyses of the flutter model, may be used to demonstrate adequate 

stability of the nominal design and failed conditions.
DATE AC 25.629-1X

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(g) Automatic Control Systems. Aeroelastic stability analyses 

of the basic configuration should include simulation of any control system for which 

interaction may exist between the sensing elements and the structural modes. Where 

structural/control system feedback is a potential problem the effects of servo-actuator 

characteristics and the effects of local deformation of the servo mount on the feedback 

sensor output should be included in the analysis. The effect of control system failures on 

the airplane aeroelastic stability characteristics should be investigated. Failures which 

significantly affect the system gain and/or phase and are not shown to be extremely 

improbable should be analysed.

b. Testing. The aeroelastic stability certification test program may consist 

of ground tests, flutter model tests, and flight flutter tests. Ground tests may be used for 

assessment of component stiffness and for determining the vibration modal characteristics 

of airplane components and the complete airframe. Flutter model testing may be used to 

establish flutter trends and validate aeroelastic stability boundaries in areas where 

unsteady aerodynamic calculations require confirmation. Full scale flight flutter testing 

provides final verification of aeroelastic stability. The results of any of these tests may be 

used to provide substantiation data, to verify and improve analytical modeling procedures 

and data, and to identify potential or previously undefined problem areas.

(1) Structural Component Tests. Stiffness tests or ground vibration 

tests of structural components are desirable to confirm analytically predicted 

characteristics and are necessary where stiffness calculations cannot accurately predict 

these characteristics. Components should be mounted so that the mounting 

characteristics are well defined or readily measurable.

(2) Control System Component Tests. When reliance is placed on 

stiffness or damping to prevent aeroelastic instability, the following control system tests 

should be conducted. If the tests are performed off the airplane the test fixtures should 

reflect local attachment flexibility.

(a) Actuators for primary flight control surfaces and flutter 

dampers should be tested with their supporting structure. These tests are to determine the 

actuator/support structure stiffness for nominal design and failure conditions considered 

in the fail-safe analysis.

(b) Flutter damper tests should be conducted to verify the 

impedance of damper and support structure. Satisfactory installed damper effectiveness 

at the potential flutter frequencies should, however, be assured. The results of these tests 

can be used to determine a suitable, in-service maintenance schedule and replacement life 

of the damper. The effects of allowable in-service free play should be measured.
DATE AC 25.629-1X

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(3) Ground Vibration Tests.

(a) Ground vibration tests (GVT) or modal response tests are 

normally conducted on the complete conforming airplane. A GVT may be used to check 

the mathematical structural model. Alternatively, the use of measured modal data alone 

in aeroelastic stability analyses, instead of analytical modal data modified to match test 

data, may be acceptable provided the accuracy and completeness of the measured modal 

data is established. Whenever structural modifications or inertia changes are made to a 

previously certified design or a GVT validated model of the basic airplane, a GVT may 

not be necessary if these changes are shown not to affect the aeroelastic stability 

characteristics.

(b) The airplane is best supported such that the suspended 

airplane rigid body modes are effectively uncoupled from the elastic modes of the 

airplane. Alternatively, a suspension method may be used that couples with the elastic 

airplane provided that the suspension can be analytically de-coupled from the airplane 

structure in the vibration analysis. The former suspension criterion is preferred for all 

ground vibration tests and is necessary in the absence of vibration analysis.

(c) The excitation method needs to have sufficient force output 

and frequency range to adequately excite all significant resonant modes. The effective 

mass and stiffness of the exciter and attachment hardware should not distort modal 

response. More than one exciter or exciter location may be necessary to insure that all

significant modes are identified. Multiple exciter input may be necessary on structures 

with significant internal damping to avoid low response levels and phase shifts at points 

on the structure distant from the point of excitation. Excitation may be sinusoidal, 

random, pseudo-random, transient, or other short duration, non stationary means. For 

small surfaces the effect of test sensor mass on response frequency should be taken into 

consideration when analyzing the test results.

(d) The minimum modal response measurement should consist 

of acceleration (or velocity) measurements and relative phasing at a sufficient number of 

points on the airplane structure to accurately describe the response or mode shapes of all 

significant structural modes. In addition, the structural damping of each mode should be 

determined.

(4) Flutter Model Tests.

(a) Dynamically similar flutter models may be tested in the 

wind tunnel to augment the flutter analysis. Flutter model testing can substantiate the 

flutter margins directly or indirectly by validating analysis data or methods. Some aspects 

of flutter analysis may require more extensive validation than others, for example control 

surface aerodynamics, T-tails and other configurations with aerodynamic interaction and 

compressibility effects. Flutter testing may additionally be useful to test configurations 

that are impractical to verify in flight test, such as fail-safe conditions or extensive store 
DATE AC 25.629-1X

17

configurations. In any such testing, the mounting of the model and the associated 

analysis should be appropriate and consistent with the study being performed.

(b) Direct substantiation of the flutter margin (clearance 

testing) implies a high degree of dynamic similitude. Such a test may be used to augment 

an analysis and show a configuration flutter free throughout the expanded design 

envelope. All the physical parameters which have been determined to be significant for 

flutter response should be appropriately scaled. These will include elastic and inertia 

properties, geometric properties and dynamic pressure. If transonic effects are important, 

the Mach number should be maintained.

(c) Validation of analysis methods is another appropriate use of 

wind tunnel flutter testing. When the validity of a method is uncertain, correlation of 

wind tunnel flutter testing results with a corresponding analysis may increase confidence 

in the use of the analytical tool for certification analysis. A methods validation test 

should simulate conditions, scaling and geometry appropriate for the intended use of the 

analytical method.

(d) Trend studies are an important use of wind tunnel flutter 

testing. Parametric studies can be used to establish trends for control system balance and 

stiffness, fuel and payload variations, structural compliances and configuration variations. 

The set of physical parameters requiring similitude may not be as extensive to study 

parametric trends as is required for clearance testing. For example, an exact match of the 

Mach number may not be required to track the effects of payload variations on a transonic 

airplane.

(5) Flight Flutter Tests.

(a) Full scale flight flutter testing of an airplane configuration 

to VDF/MDF is a necessary part of the flutter substantiation. An exception may be made 

when aerodynamic, mass, or stiffness changes to a certified airplane are minor, and 

analysis or ground tests show a negligible effect on flutter or vibration characteristics. If 

a failure, malfunction, or adverse condition is simulated during a flight test, the maximum 

speed investigated need not exceed VFC/MFC if it is shown, by correlation of the flight test 

data with other test data or analyses, that the requirements of § 25.629(b)(2) are met.

(b) Airplane configurations and control system configurations 

should be selected for flight test based on analyses and, when available, model test 

results. Sufficient test conditions should be performed to demonstrate aeroelastic 

stability throughout the entire flight envelope for the selected configurations.

(c) Flight flutter testing requires excitation sufficient to excite 

the modes shown by analysis to be the most likely to couple for flutter. Excitation 

methods may include control surface motions or internal moving mass or external 

aerodynamic exciters or flight turbulence. Use the appropriate method of excitation for 
DATE AC 25.629-1X

18

the modal response frequency being investigated. The effect of the excitation system 

itself on the airplane flutter characteristics should be determined prior to flight testing.

(d) Measurement of the response at selected locations on the 

structure should be made in order to determine the response amplitude, damping and 

frequency in the critical modes at each test airspeed. It is desirable to monitor the 

response amplitude, frequency and damping change as VDF/MDF is approached. In 

demonstrating that there is no large and rapid damping reduction as VDF/MDF is 

approached, an endeavor should be made to identify a clear trend of damping versus 

speed. If this is not possible, then sufficient test points should be undertaken to achieve a 

satisfactory level of confidence that there is no evidence of an adverse trend.

(e) An evaluation of phenomena not presently amenable to 

analyses, such as shock effects, buffet response levels, vibration levels, and control 

surface buzz, should also be made during flight testing.

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